1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with leading edge cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section with one or more rows or stages of rotor blades that react with a high temperature gas flow in order to drive the compressor and, in the case of an aero engine a fan, or in the case of an industrial gas turbine engine, an electric generator. Increasing the turbine inlet temperature will increase the efficiency of the engine. However, the highest turbine inlet temperature is dependent on the material properties and the amount of cooling provided to the parts exposed to the high temperatures. The limiting parts are the first stage airfoils which include the first stage blades and the first stage vanes.
The leading edge of an airfoil is exposed to the highest temperature gas flow which directly hits the stagnation point of the leading edge. Prior art methods of cooling the leading edge include a combination of impingement cooling of the backside wall followed by film cooling of the external wall. FIG. 1 shows a prior art turbine blade with a 3-pass aft flowing serpentine flow cooling circuit in which the first leg 11 is located adjacent to a leading edge impingement cavity 15 and is connected to it through a row of metering and impingement holes 14 formed within the rib that separates the first leg 11 from the impingement cavity 15. The serpentine circuit also includes a second leg 12 and a third leg 13 that is positioned adjacent to the trailing edge region in which a row of exit holes 16 are formed to discharge cooling air from the serpentine circuit. Spent cooling air in the impingement cavity 15 exits the blade through one or more tip cooling holes at the blade tip. Tip cooling holes are also connected to the tip turn in the serpentine circuit to discharge some of the cooling air through the blade tip than flows through the serpentine circuit. For an airfoil with low cooling flow design, especially low leading edge impingement flow design, the radial spacing for the leading edge impingement hole 14 will be larger than the impingement jet can be spread out within the inner surface of the leading edge corner. This will induce a region with low impingement cooling area within the inner surface of the leading edge corner. This will form an area with low impingement cooling within the inner surface of the leading edge corner. This will yield a hot spot in-between the impingement hole 14 and uneven cooling for the blade leading edge impingement cooling cavity 15. In addition, cross flow effect induced by the multiple hole impingement will lower the impingement heat transfer performance. In the prior art blade design of FIG. 1, some of the cooling air from the serpentine flow circuit is diverted into the leading edge impingement cavity 15 for cooling to the leading edge wall.